Gas turbine engines for aircraft often use variable area exhaust nozzles. Such engines may operate at low power in a base control mode, wherein the nozzle area is fixed. Throttle action by the pilot sets either a fuel flow rate or a engine RPM to be achieved, with pressure distribution through the engine settling out at a new value. It is known, however, that the thrust may be increased and the overall efficacy of engine operation improved by changing the area of the nozzle to an optimum condition for the new operating mode. If the nozzle closes too much, it may cause a compressor stall, while if it opens more than is necessary, over expansion within the discharge nozzle occurs.
It is accordingly known to measure the engine pressure ratio, which is the ratio of pressure leaving the gas turbine to the pressure entering the compressor and to operate nozzle to maintain this parameter. Essentially, the pressure ratio is known for the engine design which will, for any particular RPM, provide reasonable tolerence from stall with optimum thrust.
The fan, or first stage of the compressor, of an engine, is susceptible to fan damage in various situations such as the ingestion of birds, ice or other foreign objects. The initial damage may result in a stall event. In accordance with normal procedures the nozzle is opened to an increased area until recovery from the stall, and then closed down to its normal operating position. Since fan damage has occurred, it is quite possible for the engine to continue to repeatedly stall, producing unstable operation. This is possible with a fixed nozzle condition, but even more so when the engine is operating in the engine pressure ratio mode to achieve optimum thrust.
It is an object of the invention to detect and accommodate compressor fan damage, thereby effecting a proper choice of stall recovery action.